Gyroscopic navigation system and method

ABSTRACT

A compact, lightweight, cost effective, self-contained standby electronic navigation system with high signal-to-noise ratio and good dynamic stability is provided. The system includes a first sensor module for providing a plurality of rotational rate signals, a second sensor module for providing a plurality of compensation signals, and a microcontroller module for processing the rotational rate signals and the compensation signals and sending the signals to a display for displaying attitude information, directional information, and turn coordinate information on a single screen simultaneously. In one embodiment, the first sensor module includes a plurality of rotational sensors made of piezoelectric elements. The piezoelectric elements are made from a single sheet of piezoelectric material so that the elements possess uniform characteristics, and are arranged to reduce systematic drift and random noise normally present in a rotational rate sensor. The sensors can be configured on a single multi-sensor chip.

CROSS-REFERENCE TO RELATED APPLICATION(S)

This application claims benefit under 35 U.S.C. § 119(e) to U.S.application Ser. No. 10/055,186, filed Jan. 23, 2002, entitled“Gyroscopic Navigation System and Method” (Attorney Docket No. 14143),the contents of which are incorporated herein in their entirety for allpurposes.

FIELD OF THE INVENTION

The present invention relates generally to aircraft instruments, andmore particularly, to a gyroscopic navigation system for a smallaircraft's primary instruments, such as an attitude indicator (AI), adirectional gyro (DG), and a turn coordinator (TC)/slip-skid indicator.

BACKGROUND OF THE INVENTION

All aircraft, large and small, production or experimental, depend ongyroscopes for a variety of navigational data. Most aircraft utilizemechanical or spinning-mass gyros to derive information, such as headingand attitude. Often housed in a remote location, an aircraft's gyro (orgyros) feed data back to a cockpit, which are displayed in a variety ofinstruments.

The standard primary instruments in small aircraft, such as experimentalairplanes, or aircraft used in general aviation, include an attitudeindicator, a heading indicator, and a turn coordinator/slip-skidindicator. An attitude indicator, also known as an artificial horizon,shows the relationship of a nose and wings to an aircraft's horizontalplane. In addition, a heading indicator, also known as a directionalgyro, is used to counteract errors that occur in a magnetic compassduring turns, speed changes, and turbulence. A turncoordinator/slip-skid indicator measures the rate and quality of a turn.During IFR (Instrument Flight Rules) flight, these instruments,especially an attitude indicator, are pilot's means of determining anaircraft's situation.

The standard primary instruments in small aircraft currently containvacuum or electrically-driven mechanical gyroscopes. Loss of vacuum orelectrical power, especially during instrument flight, renders theseinstruments useless and can result in pilot disorientation and, attimes, fatal crashes.

There is an increasing focus within the aviation industry about thefailure of mechanical gyroscopes and the lack of proper backup systems.The FAA recently published a safety pamphlet entitled “The SilentEmergency” that emphatically states that pilots of small aircraft shouldinstall backup systems for their aircraft. In addition, the ExperimentalAircraft Association (EAA) is also actively promoting the adoption ofstandby gyroscopic navigation system.

A standby gyroscopic navigation system has been used in many commercialaircraft. Typically, in a commercial aircraft, a second set of primaryinstruments are installed independently as a standby navigation systemto ensure that there are at least one set of properly functionedgyroscopes, in case that vacuum or electrical power of the other systemfails. However, the cost of having a second set of primary instrumentsis prohibitively expensive for a small aircraft. In addition, a smallaircraft is extremely sensitive to extra weight. Further, cockpit sizeand instrument panel space are very limited in small aircraft.

Other types of standby navigation systems have been developed, forexample, Goodrich ESIS GH-3000 system. The Goodrich system is anelectronic standby instrument system. The system provides navigationinformation such as attitude, altitude, airspeed, heading, etc., and theinformation is presented on a screen display. However, the system is notuser-friendly. According to the Goodrich system's Pilot Guide, theattitude, altitude, airspeed, heading, and other navigation features arenot displayed on a LCD screen display simultaneously. The Goodrichsystem displays navigation features on different screens selected by amode selector. A pilot must change the screens to obtain differentnavigation features in flight. This increases pilot's cockpit managementload, thereby reducing overall pilot awareness, which is considereddangerous in operating a small aircraft. In addition, the Goodrichsystem is generally too heavy and too costly for a small aircraft.

One of the main parameters related to performance of an electronicinstrument system is signal-to-noise ratio. Typically, there are twotypes of noise; random noise and correlated noise. Random noise aretypically caused by characteristics of a device, such as the sensitivityof a sensor, etc. Factors that contribute to correlated noise includetemperature, vibration, electromagnetic fields, etc. Discriminationbetween signal and noise determines the performance of an electronicinstrument system.

Another parameter related to performance of an electronic instrument isthe drift of indication, often referred to as dynamic stability. Forexample, when an aircraft makes a constant angle turn and holds at thatangle, gyro instruments should indicate holding at that angle as well.However, due to the instability, gyro instruments can only hold for acertain period of time and then tend to drift back to a level position.Even though an aircraft does not normally turn and hold an angle forlonger than 3 minutes, the drift is an important parameter forperformance. Typically, existing standby instrument systems can holdabout 3-4 minutes before drifting back to a level position. It isdesirable to have a system with longer holding time.

Accordingly, there is a need for an improved electronic navigationsystem. More specifically, there is a need for a compact, lightweight,cost effective electronic navigation system with higher signal-to-noiseratio and better dynamic stability.

SUMMARY OF THE INVENTION

To solve the above and the other problems, the present inventionprovides a gyroscopic navigation system having an internal batterybackup. The system may operate as a standby gyroscopic navigation systemthat is independent of the aircraft's primary vacuum and electricalsystems and provides key navigational information that is normallyprovided by an aircraft's primary instruments, such as an attitudeindicator, directional gyro, and turn coordinator/slip-skid indicator,in the event of a catastrophic failure of the primary instruments. Thesystem may also operate as a primary gyroscopic navigation system.

In one embodiment of the present invention, a standby gyroscopicnavigation system is disclosed as an example.

Still in one embodiment, the standby gyroscopic navigation system is athree-dimensional, solid-state gyroscopic navigation system that isbattery-driven and functions independently of the aircraft's primarypower systems, both electric and vacuum. The gyroscopic navigationsystem includes a high-performance solid-state gyroscope, an on-boardsignal processing electronics, color liquid crystal display (LCD), and aback-up battery.

Further in one embodiment of the present invention, the standbygyroscopic navigation system includes a first sensor module forproviding a plurality of rotational rate signals, a second sensor modulefor providing a plurality of compensation signals, a microcontrollermodule for processing the rotational rate signals and the compensationsignals and sending the processed signals to a display to displaynavigation information such as attitude information, directionalinformation, and turn coordinate information that is normally providedby an aircraft's primary instruments, such as an attitude gyro, adirectional gyro, and a turn coordinator/slip-skid indicator.

Still in one embodiment, the standby gyroscopic navigation systemincludes a power management module for supplying power to the firstsensor module, the second sensor module, the microcontroller module, anda LCD display, etc. The power management module includes a switchcapable of switching between a primary power source and a battery powersource. Under a normal condition, the switch is switched to connect tothe primary power source, and at the mean time, the primary power sourcecharges battery of the battery power source. In the event that theprimary power source fails, the switch is automatically switched toconnect to the battery power source. The battery power source thensupplies power to the first and second sensor modules, themicrocontroller module, and the LCD display, etc.

Further in one embodiment, the first sensor module includes a set ofrotational rate sensors to provide rotational rates for pitch, roll, andyaw angles. The second sensor module includes a set of sensors toprovide acceleration, magnetic field, and temperature compensationsignals.

Yet in one embodiment, the standby gyroscopic navigation system includesa converter for converting the rotational rate signals and thecompensation signals into digital signals before sending the signals tothe microcontroller. Still in one embodiment, the provided navigationinformation is displayed on the single screen of the displaysimultaneously.

Further in one embodiment, the rotational rate sensors are solid-state,high-performance rotational rate sensors made of piezoelectric elements.The piezoelectric elements are made from a single sheet of piezoelectricmaterial so that the elements possess uniform characteristics.

In one embodiment, the piezoelectric elements are arranged andconfigured in a circular shape with an inner ring and an outer ringdisposed on a suspended membrane. The piezoelectric elements on theouter ring and the inner ring are a differential pair and areelectrically connected to the pair of piezoelectric elements on theopposite side of the circular shape, respectively. Such arrangementreduces the systematic drift and random noise normally presented in agyro rotational rate sensor. Alternatively, the piezoelectric elementsare arranged and configured in an oval shape. The resonant frequency inthe X-direction shifts relative to the resonant vibration frequency inthe Y-direction in proportion to a ratio approximately the same ratio asthe length to the width of the oval shape. The offset in resonantfrequency further enhances the stability of the system because theresonant vibration modes parallel to a surface are not closely coupled.

In an additional embodiment of the present invention, the sensors of thegyroscopic navigation system, such as the gyro and compensation sensors,are configured on a multi-sensor silicon chip. The advantage of thisembodiment is that the compensation signals are highly correlated to therotational sensor or gyro signals. The high correlation is achieved fromthe fact, but not limited to, that the sensors are physically located invery close proximity and witness the same effects, manufactured usingthe same fabrication steps, and affected in a similar manner bytemperature and electronic noise due to their similar structure. At thesystem or sub-system level, the multi-sensor chip also provides animproved degree of accuracy and compensation.

Furthermore, attitude, heading and bank angle information are displayedon a single display. One of the advantages of the present invention isthat the system is autonomous. The system is connected to an aircraft'selectrical system, but in the event of a failure, the systemautomatically reverts to battery power. The remarkable power efficiencyof solid-state gyros enables the system to run for hours on batterypower.

Another advantage of the present invention is that the system is a plugand play, maintenance free unit. The absence of moving parts lends thesolid-state gyros an exceptionally long life and, unlike traditionalgyros, the gyros of the present invention generally will not needrebuilding, thereby significantly reducing the cost.

A further advantage of the present invention is that the system islightweight. Many pilots, especially those with a smallerhigh-performance aircraft, are particularly sensitive to extra weight.The weight of the system is preferably approximately 2.0 pounds or less,more preferably 1.0 pound or less.

An additional advantage of the present invention is that the system isrugged and is designed to withstand shocks of 1000 Gs or more.

Yet another advantage of the present invention is that the system issimple to retrofit. The system is compact and capable of fitting astandard instrument panel opening in the cockpit of an aircraft.

With respect to the heart of the standby navigation system, i.e. thegyros, the system in accordance with the principles of the presentinvention includes a solid-state micro-gyroscope (or “gyro”). The gyrogenerates a voltage output proportional to rotational rate. The gyroutilizes a plurality of precision thin-film piezoelectric elements todetect rotation, such as pitch, roll, and yaw, while rejecting spuriousnoise created by vibration, thermal gradients, and electromagneticinterference. During a normal operation, selected piezoelectric elementson the gyro are driven by a periodic signal to create a controlledmechanical oscillation. When the gyro is subjected to rotational motion,such as pitch, roll, or yaw, a characteristic voltage is produced acrossother piezoelectric elements on the gyro, according to the CoriolisEffect. These voltages are amplified and filtered to extracthigh-fidelity signals proportional to the rate of rotation.

Generally, piezoelectric materials are used in a variety of sensors andactuators. Piezoelectric materials convert mechanical energy toelectrical energy and vice versa. For instance, if pressure is appliedto a piezoelectric crystal, a voltage is generated in proportion therebyproducing the function of a sensor. Generation of an electrical signalin response to an applied force or pressure is known as the “primarypiezoelectric effect”. Similarly, if an electrical voltage is applied toa piezoelectric crystal, it expands in proportion as an actuator.Geometric deformation (expansion or contraction) in response to anapplied electric field is known as the “secondary piezoelectric effect”.Whether operated as a sensor or actuator, electrically-conductiveelectrodes must be appropriately placed on a crystal for collection orapplication of the electrical energy, respectively. Therefore, apiezoelectric sensor/actuator generally includes a) a portion ofpiezoelectric material, and b) electrically-conductive electrodessuitably arranged to direct/supply electrical energy to/from anelectrical circuit, e.g. an amplifier/an external power source.

Piezoelectric materials have been utilized in the art to create avariety of simple sensors and actuators. Examples of sensors includevibration sensors, microphones, and ultrasonic sensors. Examples ofactuators include ultrasonic transmitters and linear positioningdevices. However, in most of these art examples, bulk piezoelectricmaterial is machined and assembled in a coarse manner to achievelow-complexity devices.

By contrast, the gyroscopic navigation system in accordance with theprinciples of the present invention utilizes piezoelectric materials ina thin-film format. The thin-film enables transducers with a far higherdegree of complexity of accuracy. Thin-films offer the following keyadvantages:

-   -   1) Matching—Thin-film piezoelectric materials are deposited and        defined on an atomic scale by utilizing fabrication processes        common in semiconductor industry. The result is that thin-film        piezoelectric elements can be consistently manufactured with        elements matching more than hundreds of times better than        conventional bulk machined devices.    -   2) Density—Thin-film piezoelectric elements are defined using        microlithography, a process which enables extremely small        dimensions (less than 0.001 mm, or 1 micron) to be delineated in        a consistent and controlled manner. The result is that a large        number of precision piezoelectric elements can be defined on a        single microscopic transducer device, enabling differential        arrangements that reduce noise, e.g. drift.    -   3) Accuracy—In a thin-film format, piezoelectric materials        exhibit reduced levels of random noise. At a system level, the        effect of lower noise is more accurate readings.    -   4) Low-cost—Thin-film piezoelectric elements are defined using        batch processing techniques common in the semiconductor        industry. A typical deposition, pattern transfer, and etch        sequence on a single silicon wafer defines literally millions of        precision piezoelectric elements on thousands of transducers.

The above and other advantages are inherent to the present invention andenable novel configurations and unique features that increase theoverall device and system performance.

While multiple embodiments are disclosed, still other embodiments of thepresent invention will become apparent to those skilled in the art fromthe following detailed description, wherein is shown and described onlythe embodiments of the invention, by way of illustration, of the bestmodes contemplated for carrying out the invention. As will be realized,the invention is capable of modifications in various obvious aspects,all without departing from the spirit and scope of the presentinvention. Accordingly, the drawings and detailed description are to beregarded as illustrative in nature and not restrictive.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of one embodiment of a gyroscopic navigationsystem, in accordance with the principles of the present invention,showing a LCD display of the gyroscopic navigation system.

FIG. 2 is a functional block diagram of one embodiment of the gyroscopicnavigation system, in accordance with the principles of the presentinvention.

FIG. 3 is a functional block diagram of one embodiment of sensor modulesand data signals of the gyroscopic navigation system, in accordance withthe principles of the present invention.

FIG. 4 is a cross-sectional view of one embodiment of a micro-gyro ofthe gyroscopic navigation system, in accordance with the principles ofthe present invention.

FIG. 5 is a top view of one embodiment of the micro-gyro of thegyroscopic navigation system, showing one arrangement of piezoelectricelements, in accordance with the principles of the present invention.

FIG. 6 is a schematic view of a first embodiment of a gyro electricalconfiguration for parallel connection of symmetric and differentialelements without feedback for an z-axis vibration control, in accordancewith the principles of the present invention.

FIG. 7 is a schematic view of a second embodiment of a gyro electricalconfiguration for parallel connection of symmetric and differentialelements with feedback for z-axis vibration control, in accordance withthe principles of the present invention.

FIG. 8 is a schematic view of a third embodiment of a gyro electricalconfiguration for parallel connection of symmetric and differentialelements with feedback for adaptive z-axis vibration control, inaccordance with the principles of the present invention.

FIG. 9 is an operational flow of an operation of the gyroscopicnavigation system, in accordance with the principles of the presentinvention.

FIG. 10 is a top view of the embodiment of the micro-gyro of thegyroscopic navigation system, showing an alternative arrangement ofpiezoelectric elements, in accordance with the principles of the presentinvention.

FIGS. 11(a) and 11(b) are schematic views of resonant vibrationfrequency comparison between a circular shape arrangement and an ovalshape arrangement of piezoelectric elements.

FIG. 12(a) is a cross-sectional schematic view of the gyroscopicnavigation system configured on a multi-sensor chip with shared lowerelectrode layer and shared piezoelectric layer in accordance with theprinciples of the present invention.

FIG. 12(b) is a cross-sectional schematic view of the gyroscopicnavigation system configured on a multi-sensor chip with separate lowerelectrode layer and separate piezoelectric layer in accordance with theprinciples of the present invention.

FIG. 13 is a top schematic view of one embodiment of the gyroscopicnavigation system configured on a multi-sensor chip in accordance withthe principles of the present invention.

FIG. 14 is a functional block diagram of one embodiment of thegyroscopic navigation system configured on a multi-sensor chip inaccordance with the principles of the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

As an example, a standby gyroscopic navigation system 100 is disclosedin preferred embodiments of the present invention. It is appreciatedthat the system in accordance with the present invention may be used asa primary navigation system. The preferred embodiments of the presentinvention provide a three-dimensional, solid-state gyroscopic navigationsystem that is battery-driven and functions independently of theaircraft's primary power systems, both electric and vacuum.

Preferred embodiments of the standby gyroscopic navigation system 100 asshown in FIGS. 1-10 include high-performance solid-state gyroscopes,on-board signal processing electronics, color liquid crystal display(LCD), a back-up battery, and other sensors, such as magnetometers, etc.

The present invention reflects the trend to reduce pilot workload bydisplaying multiple functions on a single display. The system 100simultaneously displays data, traditionally provided by three separateinstruments, attitude indicator (“AI”), directional gyro (“DG”) and turncoordinator (“TC”)/slip-skid indicator, on a single display. Further,these traditionally analog instruments are simulated digitally, therebyeasing the transition from traditional analog instruments to modern,all-digital system.

FIG. 1 illustrates a schematic view of one embodiment of the standbygyroscopic navigation system 100 in accordance with the principles ofthe present invention. The system 100 is a miniature, compact,self-contained, lightweight unit. The system 100 preferably includes aLCD display 102 installed in a housing 104. The display 102 providesgyroscopic navigation information that is normally provided by anaircraft's primary instruments. The navigation information includes anAl information 106, a DG information 108, and a TC/slip-skid information110. A reset push button 112 and a dial turn/push button 114 aredisposed on the housing 104 and are used for programming preflightsettings and in-flight adjustment. These gyroscopic navigationinformation are preferably presented on a single, clear, easy-to-readscreen of the LCD display 102 simultaneously. The inclusion of the threeprimary navigation instruments in a single display greatly reduces thenumber of instruments that must be scanned during instrument flight andthe uncertainty about the physical situation (e.g. attitude and heading)of the aircraft.

The DG information 108 provides a simulated analog directional “tape”with an inset digital readout 118 displaying the current directionalheading of an aircraft in flight. This configuration gives a similarreading to the standard rotating DG with the accuracy and precision of adigital device.

The AI information 106 provides digital simulation of the standardanalog Al. Both pitch (i.e. the elevation of the nose of an aircraft)and roll (i.e. the elevation of the wing of an aircraft), relative tothe horizon, are displayed in a manner consistent with analoginstruments.

The TC/slip-skid information 110 provides a digital simulation of turncoordinator/slip-skid ball, typically incorporated into an analog turncoordinator/slip-skid indicator. In an analog turn coordinator/slip-skidindicator, the mechanical slip/skid ball moves in relation to aircraftmovement about a Z-axis of an aircraft, also known as the “yaw”movement. The system of the present invention recreates this measurementin a non-mechanical, digital format that gives a pilot more accurate yawreadings at a glance, in a well-accepted visual format.

The system 100 preferably weighs less than 2.0 pounds, more preferably1.0 pound or less, and is able to withstand shocks of up to 1,000 Gs.The system 100 is preferably designed to fit into a standard instrumentpanel opening. In addition, the depth of the system 100 is preferably nomore than 2 inches. As shown, the system 100 can be securely mountedonto an aircraft panel by a set of screws 116 or other similar mountingmeans. It is appreciated to a person skilled in the art that otherappropriate sizes and weights for the system 100 can be used withoutdeparting from the scope of the invention.

The screen of the display 102 has preferably two colors to distinguishthe areas above and below the artificial horizon. In one embodiment, thecolor above the artificial horizon is blue, and the color below isbrown. It is appreciated that other color arrangements for the screencan be used within the scope of the present invention.

As shown in FIG. 1, the standby gyroscopic navigation system 100indicates that an aircraft in flight is turning towards left atapproximately 10 degrees, and the aircraft is climbing at an angle ofapproximately 3 degrees. In addition, the ball of the turncoordinator/slip-skid indicator 110 is at the middle, which indicatesthat the aircraft is coordinated in flight.

Also, a direction readout box 118 indicates that the heading of anaircraft is 360°. This allows a pilot to quickly acknowledge the headingof the aircraft without reading or counting directional markers 120.

FIG. 2 illustrates a functional block diagram of one embodiment of thestandby gyroscopic navigation system 100 in accordance with theprinciples of the present invention. The system 100 integrates severalkey modules to collect, process, and display pertinent flightinformation such as pitch, roll, yaw, signals, etc. The system 100includes a first sensor module 122 for providing a plurality ofrotational rate signals, a second sensor module 124 for providing aplurality of compensation signals, a microcontroller module 126 forprocessing the rotational rate signals and the compensation signals andsending the processed signals to a display driver 128. The displaydriver 128 processes the signals and displays gyro navigationalinformation that is normally provided by an aircraft's primaryinstruments, such as attitude gyro, directional gyro, and turncoordinator/slip-skid indicator, on the display 102.

The first sensor module 122 includes a plurality of rotational ratesensors or micro gyros 136. One of the sensors 136 is for a directionalgyro 132, and the other one is for an attitude gyro 130 (see FIG. 3). Adriver 134, e.g. an oscillator, provides oscillation signals to therotational rate sensors 136 and receives feedback from a filter 160. Therotational rate sensors 136 provide rotational rates for pitch, roll,and yaw angles of an aircraft. The output voltage signals of therotational rate sensors 136 are proportional to the rate of rotation inthe pitch, roll, and yaw directions.

The second sensor module 124 includes a plurality of compensationsensors. One of the compensation sensors is a temperature sensor 140.Another compensation sensor is a DC accelerometer 142 for measuringacceleration, vibration, or gravitational force. A third compensationsensor is a magnetometer 144 for measuring the magnitude and directionof a magnetic field. The compensation sensors provide acceleration,magnetic field, and temperature compensation signals. The output voltagesignals of the compensation sensors are proportional to low-frequencylinear (i.e. DC) acceleration, temperature, and magnetic heading. It isappreciated that the compensation sensors, such as magnetometer 144, maybe placed outside of the housing of the system 100. For example, anout-board magnetometer may be placed at the rear end of an aircraft.Similarly, it is appreciated that in FIGS. 2 and 3, other components ofthe system 100 may be placed outside of the housing of the system 100.

An analog-digital converter (ADC) 138 converts the output voltagesignals from the gyro sensors 136 and compensation sensors 140, 142, 144into digital signals and sends the information to a microcontroller 126for data processing.

Based on various sensor data, the microcontroller 126 provides theactual attitude, direction, and turn coordination information of theaircraft and sends the information to a LCD display 102. Themicrocontroller 126 performs a plurality of functions that process thesignals from the converter 138 based on suitable software 150 orfirmware instructions stored in memory 149. A CPU processor 146 of themicrocontroller 126 determines attitude and heading reference. A memory148 provides predetermined factory trim attitude information. Thememories 148, 149 can be any suitable memories, such as EPROM, etc., andcan be implemented either inside or outside of the microcontroller 126.

The standby gyroscopic navigation system 100 also includes a powermanagement module 152 for supplying power to the first sensor module122, the second sensor module 124, the microcontroller module 126, andthe display 102. The power management module 152 includes a switch 153capable of switching between a primary power source 154 and a batterypower source 156. Under a normal condition, the switch is switched toconnect to the primary power source 154, and at the mean time, theprimary power source 154 charges battery of the battery power source156. In the event that the primary power source 154 fails, the switch isswitched to connect to the battery power source 156. The battery powersource 156 then supplies power to the first and second sensor modules122, 124, the microcontroller module 126, and the display 102, etc.Accordingly, the power management module 152 provides a constant voltagesupply to the system 100, charging the battery power source 156 when theprimary power source 154 is available, and reverting to the batterypower source 156 if the primary power source 154 fails.

To accurately determine the attitude of an aircraft, the microcontroller126 relies on several sensor inputs and a compensation method. The rawdata from the sensors are relatively “clean” (i.e. signal-to-noise ratiois high) due to the differential nature of micro-gyros (see detailsbelow). Additional accuracy can be attained by removing signalperturbations caused by vibration, linear acceleration, altitude andtemperature fluctuations. Also, an outer housing may be provided in thepresent invention for electromagnetic isolation (EMI), shock resistance,thermal isolation, and protection from other environmental anomalies,such as pressure fluctuation and humidity. Additionally, an integratedmagnetometer provides accurate directional information over a longsettling time.

Data signals XA, XB, XC sent to the microcontroller 126 are shown inFIG. 3. Note that reference signals XD, XE are also sent from each gyrothat is proportional to a Z-axis output. These reference signals XD, XEare indicators of gyro operation and is analyzed by the microcontroller126 to scale the raw gyro outputs and determine if the data is valid. Abuilt-in self-test (BIST) routine may also interrogate these referencesignals to ensure that the gyro is operating correctly, and that datasignals are valid.

The microcontroller 126 operates from an operational program flow whichcontains the compensation, data scaling, and attitude extractionmethods. The overall operational flow of the systems 120 is shown laterin FIG. 9.

To some extent, each of the gyro outputs is responsive to all othersensor outputs and must be compensated to achieve highly accurate data.For instance, a pitch signal is most strongly related to an actual pitchrate of an aircraft in flight, but is also impacted by linearacceleration (or g-forces), temperature, and even the roll rate of anaircraft. Table 1 below defines each of the sensory inputs XA-XJ to themicrocontroller 126, and Table 2 defines the relationship between themwhere the coefficients, ij, are the constants of proportionality betweenany two sensor signals for a linear relationship. It is appreciated to aperson skilled in the art that other functions besides a linearrelationship can be used within the scope of the present invention, forexample, exponential, polynomial, logarithmic relationship. TABLE 1Definition of sensor data variables Variable Name Description Units XAPitch Rate Signal deg/sec XB Roll Rate Signal deg/sec XC Yaw Rate Signaldeg/sec XD Rate reference Pitch/Roll deg/sec XE Rate reference Yawdeg/sec XF Magnetic Signal X T XG Magnetic Signal Y T XH Magnetic SignalZ T XI Linear Acceleration X g XJ Linear Acceleration Y g XK LinearAcceleration Z g XL Temperature ° C.

TABLE 2 Parametric relationships True Pitch YA = α_(AA)XA + α_(AB)XB +α_(AC)XC + α_(AD)(XA/XD) + Rate YA α_(AE)XE + α_(AF)XF + α_(AG)XG +α_(AH)XH + α_(AI)XI + α_(AJ)XJ + α_(AK)XK + α_(AL)XL True Roll YB =α_(BA)XA + α_(BB)XB + α_(BC)XC + α_(BD)(XB/XD) + Rate YB α_(BE)XE +α_(BF)XF + α_(BG)XG + α_(BH)XH + α_(BI)XI + α_(BJ)XJ + α_(BK)XK +α_(BL)XL True Yaw YC = α_(CA)XA + α_(CB)XB + α_(CC)XC + α_(CD)XD + RateYC α_(CE)(XC/XE) + α_(CF)XF + α_(CG)XG + α_(CH)XH + α_(CI)XI +α_(CJ)XJ + α_(CK)XK + α_(CL)XL

As an example, for the compensated pitch rate YA (deg/sec), thecoefficients:α_(AA)=0.100, α_(AB)=−0.100, α_(AC)=−0.100, α_(AD)=0.900, α_(AE)=−0.000,α_(AF)=−0.001 deg/sec/T, α_(AG)=−0.001 deg/sec/T, α_(AH)=−0.001deg/sec/T,α_(AI)=−0.100 deg/sec/g, α_(AJ)=−0.010 deg/sec/g, α_(AK)=−0.010deg/sec/g,α_(AL)=−0.010 deg/sec/° C.

It is appreciated that other suitable factors may be used within thescope of the present invention.

Under normal operation of the system 100, an operating programpreferably runs in a continuous loop. Within each loop, the first stepis to assess the condition of the system. The data is collected so thatthe system determines whether the data falls within an expected range.If the data does not fall within the expected range, an error signal isgenerated, and the program may revert to a warning signal that alertsthe user of possible failure. If the data is valid, the next step is toperform an average of several loop values, e.g. about ten (10) loopvalues. This averaging operation tends to smooth the data and reduce anynoise generated during the analog-digital conversion. The averaged datais then subjected to the compensation method, such as the one outlinedin Table 1 and Table 2. The final compensation operation is to averageover a very long time constant, for example, approximately 10 minutes,to determine any systematic offset in the rate signal. Any residualoffset determined by this final averaging operation is then subtractedfrom the rate signal. After these compensation sequences, the threeprimary rotational rates, YA, YB, and YC, are integrated to achieve afinal attitude angle. This attitude information is sent to the LCDdisplay 102. The operation of the system is further illustrated in aflow chart shown later in FIG. 9.

Still in FIG. 3, in the first sensor module 122, a first set of signalsfrom the driver 134 are modulated with a carrier frequency by amodulator 162 and sent to the attitude gyro 130. The attitude gyro 130provides rotational rate signals for a pitch angle and a roll angle. Thesignals are then sent to a low noise amplifier (LNA) 158 to be amplifiedand to a filter 160 for reducing random electrical noise. The filteredsignals are outputted to a demodulator 164. The signals are thendemodulated by the demodulator 164 so as to filter out the carrierfrequency to obtain the signals representing pitch and roll angles. Thesignals from the filter 160 are also sent back to the modulator 162. Forillustration purposes, the gyro 130 and 132 are shown in two separatedata paths. It is appreciated that the gyros 130 and 132 can beimplemented in one data path within the scope of the present invention.

Also in FIG. 3, similar to the attitude gyro 130, signals from thedriver 134 are modulated with a carrier frequency by a modulator 162 andsent to the directional gyro 132, the directional gyro 132 providesrotational rate signals for a yaw angle. The signals are then sent tothe low noise amplifier (LNA) 158 to be amplified and the filter 160 forreducing random electrical noise. The filtered signals are outputted tothe demodulator 164. The signals are then demodulated by the demodulator164 so as to filter out the carrier frequency to obtain the signalsrepresenting a yaw angle. The signals from the filter 160 are also sentback to the modulator 162.

In the second sensor module 124, the accelerometer 142 provides linearlow-frequency acceleration readings along the x-y-z coordinates,respectively. The magnetometer 144 provides directional readings of amagnetic field along the x-y-z coordinates, respectively. Theacceleration readings and directional readings are used to compensatecorrelated noise caused by acceleration, vibration, and magnetic field,respectively. Further, the temperature sensor 140 provides signals tocompensate correlated noise caused by temperature. The compensationsignals are then sent to the converter 138 to convert the signals todigital signals before they are sent to the microcontroller 126.

It is appreciated that the rotational rate sensors are solid-state,high-performance rotational rate sensors made of piezoelectric elements.The piezoelectric elements are made from a single sheet of piezoelectricmaterial so that the elements possess uniform characteristics. In oneembodiment as shown in FIG. 4, a micro-gyro 166 of the standbygyroscopic navigation system 100 in accordance with the principles ofthe present invention is provided. The micro-gyro 166 is a solid-statedevice fabricated on silicon wafers utilizing fabrication protocolscommon in the semiconductor industry. Similar to silicon integratedcircuits (ICs), the micro-gyro 166 is built up by a series of thinfilms, typically less than or about 1 micron (0.001 mm) in thickness.

As shown in FIGS. 4 and 5, the micro-gyro 166 includes a siliconsubstrate 165 and a cylindrical silicon proof-mass 168 (often referredto as “inertial mass” or “seismic mass”) that is suspended on a toroidalthin-film membrane 170. The silicon substrate 165 is supported betweenthe thin-film membrane 170 and a base substrate 167. Clearance 169 isprovided between the proof-mass 168 and the silicon substrate 165 suchthat the micro-gyro 166 has better shock resistance.

A plurality of thin-film piezoelectric elements, for example, X1, X2,X3, X4 in FIG. 4, and X1-X4 and Y1-Y4, and Z1-Z8 in FIG. 5 are disposedon the membrane 170. The height of the proof-mass 168 is preferablyabout 500 microns, the diameter of the proof-mass 168 is preferablyabout 400 microns, while the outer diameter of the membrane 170 ispreferably about 700 microns. The membrane 170 can be made of a varietyof different materials that exhibit flexibility, resistance to fatigue,and good thermal expansion match to the surrounding silicon substrate.Preferred materials for the membrane 170 are polycrystalline silicon orsilicon nitride with a preferred typical thickness of 1 micron.

The piezoelectric elements are formed from a single layer of metal,preferably platinum about 0.1 micron of thickness, which forms a commonlower electrode layer 171 and a single layer 172 of piezoelectric thinfilm, preferably PZT about 0.5 micron of thickness. By utilizing asingle common layer for the lower electrode layer 171 and thepiezoelectric thin film 172, matching between elements and elementdensity is increased, and these factors improve the gyro's signalfidelity. Further, by using a single common electrode layer 171 and asingle layer 172 of piezoelectric thin film, it reduces the number ofelectrical connections and cost, and improves reliability of the device.The piezoelectric elements are defined by upper metal electrodes 174,preferably platinum about 0.1 micron of thickness. Since thepiezoelectric thin film 172 is non-conductive, each piezoelectricelement is defined by the upper electrode 174 alone, and electricalcommunication between elements is negligible.

FIG. 5 illustrates a top view of one embodiment of the micro-gyro 166 ofthe standby gyroscopic navigation system 100, showing one exemplaryarrangement of piezoelectric element placement, in accordance with theprinciples of the present invention. Piezoelectric elements X1-X4,Y1-Y4, and Z1-Z8 are arranged and configured in a circular shape on thesuspended membrane toroid 170 bound by an inner ring 176 and an outerring 178. The arrangement of piezoelectric elements includesdifferential pairs (i.e. X1 and X2) that reside on adjacent inner andouter regions of the membrane 170. Each pair is configured for optimalmatching thereby minimizing random noise. For example, the piezoelectricelements X1 and X2 have identical electrode area and are placed atminimum spacing. Equal area of the elements allows the device to havematched impedance because the elements with equal area have the samematerial properties as well as the same temperature-electrical response.

In addition, an identical mirror-image pair is located on opposite sideof the proof-mass 168 (i.e. X1/X2 and X3/X4). During operation, thesequad pairs (2X2) generate voltages associated with motion along aparticular coordinate axis. The differential nature and symmetricplacement along the coordinate axes allows motion in other directions tobe rejected, thereby increasing the signal accuracy. The amount of“off-axis rejection” is largely contributed by the symmetry of thepairs, matching of the elements, and precision placement. Sucharrangement reduces the systematic drift and random noise normallypresent in a rotational rate sensor, thereby dramatically improving theperformance of the system 100.

Sensing operation of the system 100 is based on the Coriolis Effect. Ina normal operation, a periodic voltage is applied to elements Z1, Z3,Z5, and Z7. By the secondary piezoelectric effect (see definition in theSummary of the Invention), the membrane 170 under these four Z-elementsis deflected, and the proof-mass 168 is driven into vibration along theZ-axis (perpendicular to the surface in FIG. 5) at the same periodicrate as the applied voltage. If a rotation is applied around the X-axis,a Coriolis force forms in the Y-axis direction. Similarly, if a rotationis applied around the Y-axis, a Coriolis force forms in the X-axisdirection. The Coriolis force is proportional to the weight of theproof-mass 168, the oscillation frequency, the magnitude of oscillation,and the rate of rotation. The piezoelectric elements X1, X2, X3, and X4detect the Coriolis force along the X-axis that is associated withrotation about the Y-axis. Similarly, the piezoelectric elements Y1, Y2,Y3, and Y4 detect the Coriolis force along the Y-axis that is associatedwith rotation about the X-axis.

The elements Z2, Z4, Z6, and Z8 can be used for several differentfunctions. In one embodiment as shown in FIG. 6, these elements may bedriven with a periodic voltage 180° out of phase with respect toelements Z1, Z3, Z5, and Z7. Output voltages Vo_(p), VO_(r) (“p” standsfor “pitch”, and “r” stands for “roll”) are sent to the filter 160. In apreferred embodiment, elements Z2, Z4, Z6, and Z8 are operated assensors that provide an output voltage proportional to the vibrationalong the Z-axis as shown in FIG. 7. In still another embodiment, thevibration along the Z-axis can be tuned and equalized by independentlysensing and driving each of the Z-elements as shown in FIG. 8 to furtherreduce non-idealities in the vibration mode. In the embodiments of FIG.7 and FIG. 8, output voltages Vo_(ref), Vo_(ref1), Vo_(ref2), Vo_(ref3),Vo_(ref4) generated from the Z2, Z4, Z6, and Z8 elements can be used ina feedback loop to control the periodic driving signal. This feedbackmode of operation provides immunity to temperature variation and otherenvironmental anomalies and improves the overall fidelity of the gyroperformance. The quality of piezoelectric matching and symmetry enablethis mode of operation.

It is appreciated that the micro gyro 166 can be used in a variety ofapplicable industries for detecting sensitive motion movements inaccordance with the principles of the present invention. For example,the micro gyro 166 can be used in the automobile industry for detectingroll-over motion, anti-lock brake motion, etc.; in the consumerelectronics industry for motion activating Personal Data Assistant (PDA)devices, cell phones, etc., for remote data entry, etc., or forcompensating lens based on camera motion; in the medical device industryfor monitoring patients' activities, bio-response, automaticallyadjusting devices, such as defibrillators, pacemakers, etc.; or in thedefense industry for detecting motion of satellite dishes, etc. Theapplication of the micro gyro 166 on the gyroscopic navigation system100 for aircraft instruments is merely one example of many applications.

FIG. 9 illustrates an operational flow diagram of a standby gyroscopicnavigation operation 180 in accordance with the principles of thepresent invention. The operation 180 starts with an operation 182 ofpowering up the system, and an operation 184 of checking the system. Ifthe system is not operating in a normal condition from a determinationoperation 186, the system displays an error in an operation 188. If thesystem is operating in a normal condition from the determinationoperation 186, the system is reset to clear any prior attitudeinformation in an operation 190. Then, information on the display isupdated in an operation 192. Next, all sensor data are collected in anoperation 194. If the sensed data is not valid from a determinationoperation 196, the system displays error in the operation 188. If thesensed data is valid from the determination operation 196, average datafor XA to XL over the previous period of time, e.g. 1 ms (millisecond),are calculated in an operation 198.

Then, the system compensates for cross-coupling terms as shown in Table1, in an operation 200. For example, cross-coupling terms for YA is XBto XL. Next, average data for YA to YC for a period of time, such as 10minutes, are calculated in an operation 202. Then, the system integratesrate to obtain attitude information including pitch, roll, and yaw in anoperation 204. Next, a system check operation 206 is performed. If thesystem operates in a normal condition from a determination operation208, the system displays the updated attitude information in theoperation 192. If the system does not operate in a normal condition fromthe determination operation 208, the system display error in theoperation 188.

These operations continue until an aircraft stops its operation. Oncethe system displays error, the system can be restarted or ended in anoperation 209. It is appreciated that a longer period of time can becalculated in the operation 202. Accordingly, a longer holding time,when an aircraft makes a constant angle turn and holds at that angle,can be achieved. Thus, the dynamic stability is significantly improved.

FIG. 10 illustrates an alternative arrangement of piezoelectric elementsof a micro-gyro 166′ of the standby gyroscopic navigation system 100.Instead of being arranged and configured in a circular shape,piezoelectric elements X1′-X4′, Y1′-Y4′, and Z1′-X8′ are arranged andconfigured in an oval shape on the suspended membrane toroid 170 boundby an inner ring 176′ and an outer ring 178′.

Similar to the arrangement shown in FIG. 5, the arrangement in FIG. 10includes differential pairs of the piezoelectric elements that reside onadjacent inner and outer regions of the membrane 170. Each pair isconfigured for optimal matching there by minimizing random noise. Also,the pairs have identical electrode area and are placed at minimumspacing.

Further, an identical mirror-image pair is located on opposite side of aproof mass 168′. The electrodes and their respective f unctions areidentical to the circular embodiment. However, the resonant vibrationfrequency in the X-direction shifts relative to the resonant vibrationfrequency in the Y-direction in proportion to the ratio C:A as shown inFIG. 10. C is the length of the oval shape, and A is the width of theoval shape. FIG. 11(a) illustrates the resonant behavior for thecircular arrangement, and FIG. 11(b) illustrates the resonant behaviorfor the oval or elliptical arrangement. Frequency offset Δf isproportion to the ratio C:A. Such offset in resonant frequency furtherenhances the stability of the system 100 because the resonant vibrationmodes parallel to the element surface are not closely coupled.

In addition, the gyroscopic navigation system described above can beconfigured on a multi-sensor chip in accordance with the principles ofthe present invention. The gyro and the other compensation sensors, suchas the acceleration sensor, are configured and fabricated in a singlesilicon substrate. The compensation sensor is generally used to increasethe overall accuracy. The advantage of this embodiment is that thecompensation signals are highly correlated to the rotational sensorsignals. The high correlation is achieved from the fact, but not limitedto, that the sensors are physically located in very close proximity andwitness the same effects, manufactured using the same fabrication steps,and affected in a similar manner by temperature and electronic noise dueto their similar structure. At the system or sub-system level, themulti-sensor chip also provides an improved degree of accuracy andcompensation.

FIGS. 12(a) and 12(b) illustrate two different embodiments of amulti-sensor chip 210. FIG. 12(a) shows the gyroscopic navigation systemconfigured on the multi-sensor chip 210 having two sensors 216, 218 withshared lower electrode layer 212 and shared piezoelectric layer 214.FIG. 12(b) shows the gyroscopic navigation system configured on themulti-sensor chip 210 having the two sensors 216, 218 with separatelower electrode layer 212 and separate piezoelectric layer 214. It isappreciated that the number of sensors configured on the multi-sensorchip 210 can be varied within the scope of the present invention.

FIG. 13 is a top view of the multi-sensor chip 210 including the sensors216 and 218 and a plurality of electronic bonding pads 220, 222 foroutput connections. As illustrated, the sensors 216 and 218 are in acircular shape. It is appreciated that the sensors may be configured inmany other shapes, such as the oval shape as illustrated in FIG. 10.

FIG. 14 is a functional block diagram of one embodiment of thegyroscopic navigation system configured on a multi-sensor chip 210. Thefirst sensor 216 is connected to an oscillatory feedback circuit asdescribed above in FIGS. 2 and 3 to operate as the gyroscope orrotational rate sensor 122. The second sensor 218 is connected to anamplification circuit and provides acceleration outputs as describedabove in FIGS. 2 and 3 to operate as the compensation sensor, such asthe DC accelerometer 142.

Although the present invention has been described with reference topreferred embodiments, persons skilled in the art will recognize thatchanges may be made in form and detail without departing from the spiritand scope of the invention. For example, it is appreciated that thegyros may be other types of solid-state gyros, such as quartzsolid-state gyros.

1-40. (canceled)
 41. A solid-state gyro for displaying gyroscopicnavigational attitude information, direction information, and turncoordinate information simultaneously on a display, comprising: asubstrate having a proof-mass; a membrane, the proof-mass beingsuspended on the membrane; a common electrode layer being disposed onthe membrane; a sheet of piezoelectric material in a thin-film formatbeing disposed on the common electrode layer; and a plurality ofelectrodes being disposed on the sheet of piezoelectric material, therotational rate signals being outputted through the electrodes, whereineach of the electrodes, the piezoelectric material, and the commonelectrode layer form a plurality of piezoelectric elements.
 42. The gyroof claim 41, wherein the piezoelectric elements are arranged -andconfigured in a circular shape with a plurality of pairs ofpiezoelectric elements, one element in a pair is disposed on an innerring of the circular shape, and the other element in the pair isdisposed on an outer ring of the circular shape.
 43. The gyro of claim42, wherein the two elements of the pair have equal area.
 44. The gyroof claim 43, wherein each pair of piezoelectric elements has a mirrorimage pair of piezoelectric elements disposed on opposite side of anaxis passing through a center of the proof-mass.
 45. The gyro of claim41, wherein the piezoelectric elements are arranged and configured in anoval shape with a plurality of pairs of piezoelectric elements, oneelement in a pair is disposed on an inner ring of the oval shape, andthe other element in the pair is disposed on an outer ring of the ovalshape.
 46. The gyro of claim 45, wherein the two elements of the pairhave equal area.
 47. The gyro of claim 45, wherein each pair ofpiezoelectric elements has a mirror image pair of piezoelectric elementsdisposed on opposite side of an axis passing through a center of theproof-mass.
 48. An aircraft instrument system, comprising: a pluralityof aircraft primary instruments including a mechanical attitude gyro, amechanical directional gyro, a mechanical-electrical turncoordinator/slip-skid indicator; a standby gyroscopic navigation systemconnected independently of the primary instruments; electrical powerhaving a primary power source and a battery power source, the primarypower source supplying power to the primary instruments and the standbygyroscopic navigation system, the standby battery power source supplyingpower to the standby gyroscopic navigation system to provide attitudeinformation, directional information, and turn coordination informationwhen the primary power source fails; and wherein the standby gyroscopicnavigation system includes a solid-state gyro for displaying gyroscopicthe attitude information, the directional information, and the turncoordinate information simultaneously on a display.
 49. The system ofclaim 48, wherein the solid-state gyro comprises: a substrate having aproof-mass; a membrane, the proof-mass being suspended on the membrane;a common electrode layer being disposed on the membrane; a sheet ofpiezoelectric material in a thin-film format being disposed on thecommon electrode layer; and a plurality of electrodes being disposed onthe sheet of piezoelectric material, the rotational rate signals beingoutputted through the electrodes, wherein each of the electrodes, thepiezoelectric material, and the common electrode layer form a pluralityof piezoelectric elements.
 50. The system of claim 49, wherein thepiezoelectric elements are arranged and configured in a circular shapewith a plurality of pairs of piezoelectric elements, one element in apair is disposed on an inner ring of the circular shape, and the otherelement in the pair is disposed on an outer ring of the circular shape.51. The system of claim 50, wherein the two elements of the pair haveequal area.
 52. The system of claim 51, wherein each pair ofpiezoelectric elements has a mirror image pair of piezoelectric elementsdisposed on opposite side of an axis passing through a center of theproof-mass.
 53. The system of claim 48, wherein the piezoelectricelements are arranged and configured in an oval shape with a pluralityof pairs of piezoelectric elements, one element in a pair is disposed onan inner ring of the oval shape, and the other element in the pair isdisposed on an outer ring of the oval shape.
 54. The system of claim 53,wherein the two elements of the pair have equal area.
 55. The system ofclaim 53, wherein each pair of piezoelectric elements has a mirror imagepair of piezoelectric elements disposed on opposite side of an axispassing through a center of the proof-mass.